Gas turbine engine fan drive gear system damper

ABSTRACT

A gas turbine engine includes a fan section, a turbine section coupled to the fan section through a geared architecture, a first member supporting the geared architecture, and a second member supporting the first member relative to an engine static structure. The first member and the second members are spaced apart from one another at a location. A damper includes opposing ends and extending between the first member and the second member at the location for limiting relative movement between the first member and the second member. A fan drive gear system and a method are also disclosed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.14/036,927, filed on Sep. 25, 2013, which is a continuation of U.S.patent application Ser. No. 13/557,515, filed Jul. 25, 2012, now U.S.Pat. No. 8,573,926 granted Nov. 5, 2013, which is a continuation of U.S.patent application Ser. No. 13/432,699 filed Mar. 28, 2012, now U.S.Pat. No. 8,529,197 granted Sep. 10, 2013.

BACKGROUND

This disclosure relates to a damper for a fan drive gear system for agas turbine engine.

Gear trains are used in gas turbine engines to provide a gear reductionbetween a turbine section and a fan, for example. The gear train issupported relative to a static structure. During operation, the geartrain generates vibrational inputs to the static structure and othercomponents, which may be undesirable. Additionally, the supportingstructure may transmit vibrational inputs to the fan drive gear systemthat may be coincident or undesirable to the fan drive gear system.Typically, a flex support having a bellow secures the gear train to thestatic structure to permit some relative movement between the gear trainand the static structure.

SUMMARY

A gas turbine engine according to an exemplary embodiment of thisdisclosure, among other possible things includes a fan section, aturbine section coupled to the fan section through a gearedarchitecture, a first member supporting the geared architecture, and asecond member supporting the first member relative to an engine staticstructure. The first member and the second members are spaced apart fromone another at a location. A damper includes opposing ends and extendingbetween the first member and the second member at the location forlimiting relative movement between the first member and the secondmember.

In a further embodiment of any of the foregoing gas turbine engines, thedamper includes a seal on each of the opposing ends with the seal oneach end engaged to a corresponding one of the first member and thesecond member.

In a further embodiment of any of the foregoing gas turbine engines, thedamper includes a tube having opposing ends each with an annular groovefor receiving the seal.

In a further embodiment of any of the foregoing gas turbine engines, thefirst member is a torque frame and the second member is a flex supporthaving a bellow, the flex support grounded to a static structure.

In a further embodiment of any of the foregoing gas turbine engines, thetorque frame and flex support are secured to one another in an areaspaced radially inward from the location.

In a further embodiment of any of the foregoing gas turbine engines, thetorque frame and the flex support include first and second aperturesdisposed at the location that are aligned axially and that receive acorresponding one of the opposing ends of the damper.

In a further embodiment of any of the foregoing gas turbine engines, atleast one of the torque frame and the flex support is configured togenerate a vibrational input during operation and the damper isconfigured to damp the vibrational input.

In a further embodiment of any of the foregoing gas turbine engines,each opposing end includes annular tapers extending radially inward awayfrom the respective end's annular groove and configured to permitarticulation of the damper relative to the torque frame and the flexsupport.

In a further embodiment of any of the foregoing gas turbine engines, thefirst and second apertures and the damper are aligned parallel to anengine longitudinal axis for damping vibrational input in an axialdirection.

In a further embodiment of any of the foregoing gas turbine engines, theflex support has a bellow attached to the torque frame and multipledampers are arranged circumferentially between the torque frame and theflex support.

In a further embodiment of any of the foregoing gas turbine engines, thetorque frame supports a carrier to which star gears are mounted, a sungear is arranged centrally relative to and intermeshing with the stargears, and a ring gear circumscribes and intermeshes with the stargears.

In a further embodiment of any of the foregoing gas turbine engines,includes a compressor section including a first compressor and a secondcompressor and a combustor in communication with the compressor section.The turbine section includes a fan drive turbine coupled to drive thefan section through the geared architecture and a second turbine forwardof the fan drive turbine.

In a further embodiment of any of the foregoing gas turbine engines, thegeared architecture is configured to provide a speed reduction betweenthe fan section and the fan drive turbine greater than about 2.3:1.

In a further embodiment of any of the foregoing gas turbine engines, thegas turbine engine is a high bypass geared aircraft engine having abypass ratio of greater than about six (6).

In a further embodiment of any of the foregoing gas turbine engines, thebypass ratio is greater than about ten (10).

In a further embodiment of any of the foregoing gas turbine engines, thegas turbine engine includes a Fan Pressure Ratio of less than about1.45.

In a further embodiment of any of the foregoing gas turbine engines, afan tip speed is less than about 1150 ft/second.

In a further embodiment of any of the foregoing gas turbine engines, thefan drive turbine has a pressure ratio that is greater than about 5.

In a further embodiment of any of the foregoing gas turbine engines, thesecond turbine includes at least two rotational turbine blade stages.

In a further embodiment of any of the foregoing gas turbine engines, thefan drive turbine rotates in a direction opposite the second turbine.

A fan drive gear system according to an exemplary embodiment of thisdisclosure, among other possible things includes a first membersupporting a geared architecture, and a second member supporting thefirst member relative to a static structure. The first member and thesecond members are spaced apart from one another at a location. A damperincludes opposing ends and extending between the first member and thesecond member at the location for limiting relative movement between thefirst member and the second member.

In a further embodiment of any of the foregoing fan drive gear systems,the damper includes a seal on each of the opposing ends with the seal oneach end engaged to a corresponding one of the first member and thesecond member.

In a further embodiment of any of the foregoing fan drive gear systems,the damper includes a tube having opposing ends each with an annulargroove for receiving the seal.

In a further embodiment of any of the foregoing fan drive gear systems,the first member is a torque frame and the second member is a flexsupport having a bellow, the flex support grounded to a staticstructure.

In a further embodiment of any of the foregoing fan drive gear systems,at least one of the torque frame and the flex support is configured togenerate a vibrational input during operation and the damper isconfigured to damp the vibrational input.

A method of designing a gas turbine engine according to an exemplaryembodiment of this disclosure, among other possible things includesdefining a first member for supporting a geared architecture,configuring a second member supporting the first member relative to anengine static structure, the first member and the second members areconfigured to be spaced apart from one another at a location, andconfiguring a damper including opposing ends that extend between thefirst member and the second member at the location for limiting relativemovement between the first member and the second member.

In a further embodiment of any of the foregoing methods, includesdefining the damper to include a seal on each of the opposing ends withthe seal on each end engaged to a corresponding one of the first memberand the second member.

In a further embodiment of any of the foregoing methods, includesconfiguring the damper to include a tube having opposing ends each withan annular groove for receiving the seal.

In a further embodiment of any of the foregoing methods, includesconfiguring the first member as a torque frame and the second member asa flex support having a bellow. The flex support is configured forgrounding to the static structure.

In a further embodiment of any of the foregoing methods, includesdefining the torque frame and the flex support to include first andsecond apertures disposed at the location that are aligned axially andconfiguring the first and second apertures to receive a correspondingone of the opposing ends of the damper.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine embodiment.

FIG. 2 is a schematic view of an epicyclic gear train embodiment for afan drive gear system.

FIG. 3 is a partial cross-sectional schematic view of a fan drive gearsystem embodiment.

FIG. 4 is an enlarged view of a portion of the fan drive gear systemshown in FIG. 2.

FIG. 5 is a perspective view of a damper embodiment shown in FIG. 4.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath B whilethe compressor section 24 drives air along a core flowpath C forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure (or first) compressor section 44and a low pressure (or first) turbine section 46. The inner shaft 40 isconnected to the fan 42 through a geared architecture 48 to drive thefan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a high pressure(or second) compressor section 52 and high pressure (or second) turbinesection 54. A combustor 56 is arranged between the high pressurecompressor 52 and the high pressure turbine 54. A mid-turbine frame 57of the engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The mid-turbineframe 57 supports one or more bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis A,which is collinear with their longitudinal axes. As used herein, a “highpressure” compressor or turbine experiences a higher pressure than acorresponding “low pressure” compressor or turbine.

The core airflow C is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a star gear systemor other gear system, with a gear reduction ratio of greater than about2.3 and the low pressure turbine 46 has a pressure ratio that is greaterthan about 5. In one disclosed embodiment, the engine 20 bypass ratio isgreater than about ten (10:1), the fan diameter is significantly largerthan that of the low pressure compressor 44, and the low pressureturbine 46 has a pressure ratio that is greater than about 5:1. Lowpressure turbine 46 pressure ratio is pressure measured prior to inletof low pressure turbine 46 as related to the pressure at the outlet ofthe low pressure turbine 46 prior to an exhaust nozzle. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned per hour divided by lbf of thrustthe engine produces at that minimum point. “Fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient degR)/518.7)^0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

An example geared architecture 48 is schematically shown in a gearcompartment 96 in FIGS. 2 and 3. The geared architecture 48 includes asun gear 60, which is coupled to the inner shaft 40, as illustrated inFIG. 3. Star gears 62 are arranged circumferentially about the sun gear60 and intermesh with the sun gear 60 and a ring gear 64, whichcircumscribes the star gear 62. In one example, the ring gear 64 iscoupled to the fan 42. It should be understood that the gearedarchitecture 48 illustrated in FIGS. 2 and 3 is exemplary only and canbe configured other than illustrated.

A carrier 66 supports the star gears 62 relative to the sun gear 60 andring gear 64. A torque frame 68 is connected to the carrier 66 by pins70. The torque frame 68 is secured to the static structure 36 by a flexsupport 72, which has a bellow for permitting slight movement of thegeared architecture 48 relative to the static structure 36. In theexample, fasteners 73 secure the torque frame 68 and the flex support72, which are metallic in one example, to one another to facilitateassembly and disassembly of the geared architecture 48. However, thetorque frame 68 and flex support 72 are also spaced apart from oneanother in an axial direction at a location radially outward from thefasteners 73.

Referring to FIGS. 3 and 4, the torque frame 68 and flex support 72respectively include first and second apertures 74, 76 that are alignedwith one another in the axial direction. A damper 78, which is metallicin one example, is provided between the torque frame 68 and flex support72 and received within the first and second apertures 74, 76, the gearedarchitecture 48, provide desired stiffness and/or avoid naturalfrequencies. In one example, multiple dampers are arrangedcircumferentially between the torque frame 68 and flex support 72, asillustrated in FIG. 2. It should be understood that the dampers 78 maybe configured in any desirable configuration and more or fewer dampers78 may be used than illustrated.

Referring to FIGS. 4 and 5, the damper 78 is provided by a tube 79includes opposing ends 80 with a neck 82 arranged between the ends 80.The neck 82 has a diameter that is smaller than a diameter of the ends80. Each end 80 includes an annular groove 84 that receives a seal 92.Lateral sides 86 are provided on each end 80 with the annular groove 84arranged between the lateral sides 86. The lateral sides 86 provide anannular taper that extends radially inward from the annular groove 84.The smaller diameter neck 82 and the tapered lateral sides 86 enablesthe damper 78 to articulate within the first and second apertures 74, 76about the seals 92 during vibrations without permitting metal-to-metalcontact between the damper 78 and the torque frame 68 or the flexsupport 72.

The damper 78 includes a cavity 88 that extends along its axial lengthbetween openings 90 provided at each end 80. The cavity 88 provides aviscous damping chamber. One or more orifices 94 are provided in theneck 82, for example, and are in fluid communication with the cavity 88.The orifices 94 permit an oil-mist in the gear compartment 96 to enterthe cavity 88. Any oil collecting in the cavity 88 may drain through theorifices 94. The volume of the cavity 88 and the size, number andconfiguration of the orifices 94 are configured to damp a vibrationalinput from the geared architecture 48.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A gas turbine engine comprising: a fan section; aturbine section coupled to the fan section through a gearedarchitecture; a first member supporting the geared architecture; asecond member supporting the first member relative to an engine staticstructure, wherein the first member and the second members are spacedapart from one another at a gap; and a damper bridging the gap andincluding opposing ends and engaging the first member and the secondmember for limiting relative movement between the first member and thesecond member.
 2. The gas turbine engine as recited in claim 1, whereinthe damper includes a seal on each of the opposing ends with the seal oneach end engaged to a corresponding one of the first member and thesecond member.
 3. The gas turbine engine as recited in claim 2, whereinthe damper comprises a tube having opposing ends each with an annulargroove for receiving the seal.
 4. The gas turbine engine as recited inclaim 2, wherein the first member is a torque frame and the secondmember is a flex support having a bellow, the flex support grounded to astatic structure.
 5. The gas turbine engine as recited in claim 4,wherein the gap and damper are provided at a location, and the torqueframe and flex support are secured to one another in an area spacedradially inward from the location.
 6. The gas turbine engine as recitedin claim 5, wherein the torque frame and the flex support include firstand second apertures disposed at the location that are aligned axiallyand that receive a corresponding one of the opposing ends of the damper.7. The gas turbine engine as recited in claim 6, wherein at least one ofthe torque frame and the flex support is configured to generate avibrational input during operation and the damper is configured to dampthe vibrational input.
 8. The gas turbine engine as recited in claim 6,wherein the first and second apertures and the damper are alignedparallel to an engine longitudinal axis for damping vibrational input inan axial direction.
 9. The gas turbine engine as recited in claim 4,wherein the flex support has a bellow attached to the torque frame andmultiple dampers are arranged circumferentially between the torque frameand the flex support, and the gap is provided in an axial direction. 10.The gas turbine engine as recited in claim 9, wherein the torque framesupports a carrier to which star gears are mounted, a sun gear isarranged centrally relative to and intermeshing with the star gears, anda ring gear circumscribes and intermeshes with the star gears.
 11. Thegas turbine engine as recited in claim 10, including a compressorsection comprising a first compressor and a second compressor and acombustor in communication with the compressor section, wherein theturbine section includes a fan drive turbine coupled to drive the fansection through the geared architecture and a second turbine forward ofthe fan drive turbine.
 12. The gas turbine engine as recited in claim11, wherein the geared architecture is configured to provide a speedreduction between the fan section and the fan drive turbine greater thanabout 2.3:1.
 13. The gas turbine engine as recited in claim 12, whereinthe gas turbine engine is a high bypass geared aircraft engine having abypass ratio of greater than about six (6).
 14. The gas turbine engineas recited in claim 13, wherein the bypass ratio is greater than aboutten (10).
 15. The gas turbine engine as recited in claim 14, wherein thegas turbine engine includes a Fan Pressure Ratio of less than about1.45.
 16. The gas turbine engine as recited in claim 15, wherein a fantip speed is less than about 1150 ft/second.
 17. The gas turbine engineas recited in claim 16, wherein the fan drive turbine has a pressureratio that is greater than about
 5. 18. The gas turbine engine asrecited in claim 17, wherein the second turbine includes at least tworotational turbine blade stages.
 19. The gas turbine engine as recitedin claim 18, wherein the fan drive turbine rotates in a directionopposite the second turbine.
 20. A fan drive gear system comprising; afirst member supporting a geared architecture; a second membersupporting the first member relative to a static structure, wherein thefirst member and the second members are spaced apart from one another toprovide a gap; and a damper bridging the gap and including opposing endsand engaging the first member and the second member for limitingrelative movement between the first member and the second member. 21.The fan drive gear system as recited in claim 20, wherein the damperincludes a seal on each of the opposing ends with the seal on each endengaged to a corresponding one of the first member and the secondmember.
 22. The fan drive gear system as recited in claim 21, whereinthe damper comprises a tube having opposing ends each with an annulargroove for receiving the seal.
 23. The fan drive gear system as recitedin claim 21, wherein the first member is a torque frame and the secondmember is a flex support having a bellow, the flex support grounded to astatic structure.
 24. The fan drive gear system as recited in claim 23,wherein at least one of the torque frame and the flex support isconfigured to generate a vibrational input during operation and thedamper is configured to damp the vibrational input.
 25. A method ofdesigning a gas turbine engine comprising: defining a first member forsupporting a geared architecture; configuring a second member supportingthe first member relative to an engine static structure, wherein thefirst member and the second members are configured to be spaced apartfrom one another to provide a gap; and configuring a damper bridging thegap and including opposing ends that engage the first member and thesecond member for limiting relative movement between the first memberand the second member.
 26. The method as recited in claim 25, includingdefining the damper to include a seal on each of the opposing ends withthe seal on each end engaged to a corresponding one of the first memberand the second member.
 27. The method as recited in claim 26, includingconfiguring the damper to include a tube having opposing ends each withan annular groove for receiving the seal.
 28. The method as recited inclaim 26, including configuring the first member as a torque frame andthe second member as a flex support having a bellow, wherein the flexsupport is configured for grounding to the static structure.
 29. Themethod as recited in claim 28, including defining the torque frame andthe flex support to include first and second apertures disposed that arealigned axially and configuring the first and second apertures toreceive a corresponding one of the opposing ends of the damper.
 30. Agas turbine engine comprising: a fan section; a turbine section coupledto the fan section through a geared architecture; a first membersupporting the geared architecture; a second member supporting the firstmember relative to an engine static structure, wherein the first memberand the second members are spaced apart from one another at a location;and a damper including opposing ends and extending between the firstmember and the second member at the location for limiting relativemovement between the first member and the second member, wherein thedamper comprises a tube having opposing ends each with an annular groovefor receiving a seal, wherein each opposing end includes annular tapersextending radially inward away from the respective end's annular grooveand configured to permit articulation of the damper relative to thefirst and second members.